This invention relates generally to methods for selectively coating internal passageways of an object with protective coatings having different thicknesses and to objects having such selectively coated internal passageways. The invention has particular use when the object being coated or which is so coated is a gas turbine blade, but the invention is not limited to gas turbine blades.
In an aircraft gas turbine (jet) engine, air is drawn into the front of the engine, compressed by a shaft-mounted compressor, and mixed with fuel. The mixture is burned, and the hot combustion gases are passed through a turbine mounted on the same shaft. The flow of combustion gas turns the turbine by impingement against an airfoil section of the turbine blades and vanes, which turns the shaft and provides power to the compressor. The hot exhaust gases flow from the back of the engine, driving it and the aircraft forward.
The hotter the combustion and exhaust gases, the more efficient is the operation of the jet engine. There is thus an incentive to raise the combustion and exhaust gas temperatures. The maximum temperature of the combustion gases is normally limited by the materials used to fabricate the hot-section components of the engine. These components include the turbine vanes and turbine blades of the gas turbine, upon which the hot combustion gases directly impinge. In current engines, the turbine vanes and blades are made of nickel-based superalloys, and can operate at temperatures of up to approximately 980–1150 degrees Celsius, or roughly 1800–2100 degrees Fahrenheit. These components are subject to damage by oxidation and corrosive agents.
Many approaches have been used to increase the operating temperature limits and service lives of the turbine blades and vanes to their current levels while achieving acceptable oxidation and corrosion resistance. The composition and processing of the base materials themselves have been improved. Cooling techniques are used, as for example by providing the component with internal cooling passages through which cooling air is flowed. However, as engine temperatures increase, the temperature of available cooling air also increases.
In at least one known configuration of gas turbine blade, a portion of the outer surfaces of the turbine blades is coated with a protective coating. One type of protective coating includes an aluminum-containing protective coating deposited upon the substrate material to be protected. The exposed surface of the aluminum-containing protective coating oxidizes to produce an aluminum oxide protective layer that protects the underlying surface.
Different portions of the outer surface of gas turbine blade require different types and thicknesses of protective coatings, and some portions require that there be no coating thereon. One known method for selective protection of the outer surfaces of a gas turbine blade is disclosed in U.S. Pat. No. 6,652,914 B1, issued Nov. 25, 2003 to Langley, et al. and assigned to General Electric Aviation Service Operation Pte. Ltd. In this method, a gas turbine blade that has previously been in service is protected by cleaning the gas turbine blade and then first depositing a precious metal layer over portions of the blade. The method includes a first deposition step in which a precious metal such as platinum is deposited on a surface of the blade, preferably by electrodeposition. The first layer is deposited on an airfoil first layer region of the airfoil. In the usual case, the first layer includes only portions of the surface of the airfoil, but not the trailing edge of the airfoil or the surface of the dovetail. The thickness of the first platinum layer is controlled to be about 0.002 mm to about 0.0032 mm, or about 0.00008 to about 0.000125 inches. In a second deposition step, a precious metal second layer is deposited overlying at least part of the platform portion of the second layer, but not overlying the airfoil portion of the first layer. The result is that the total thickness of the precious metal on the bottom side of the platform is greater than the total thickness on the airfoil.
A platinum alunimide protective coating is then formed by depositing an aluminum-containing layer overlying both the platform and the airfoil and interdiffusing the platinum and the aluminum. A vapor-phase aluminiding process is used in which baskets of chromium-aluminum alloy pellets are positioned within about 25 mm (one inch) of the gas turbine blade to be vapor-phase aluminided, in a retort. The retort containing the baskets and the turbine blade (or a plurality of blades together) are heated in an argon atmosphere at a heating rate of about 28 degrees Celsius (50 degrees Fahrenheit) per minute to a temperature of about 1080 degrees +/−14 degrees Celsius (1975 +/−25 degrees Fahrenheit), held at that temperature for about 3 hours +/−15 minutes, during which time aluminum is deposited, and then slow cooled to about 120 degrees Celsius (250 degrees Fahrenheit), and thence to room temperature. The times and temperatures may be varied to alter the thickness of the aluminum containing layer. The first, second, and third layers interdiffuse to form an interdiffused airfoil platinum aluminide protective coating over the airfoil first layer region, and a platform interdiffused platinum aluminide protective layer over the platform first layer region. A further heating can be applied to further interdiffuse the layers, and the layers cleaned. The resulting platform interdiffused protective layer has a different thickness than the airfoil interdiffused protective layer, largely as a result of differences in the thicknesses of the separately applied precious metal layers.
As noted above, however, modern gas turbine blades are cooled by passing cooling air through internal cooling passages. As engine temperatures increase, the temperature of available cooling air also increases, and corrosion can occur in these internal passages as well as on the external surfaces.
Internal coating thickness requirements for turbine blades vary depending upon location. For example, a thin coating is required in high stress areas such as the blade shank, and a robust, thick coating is required in other areas such as airfoil cavities to protect against the environment. If only a single thickness can be accomplished, the areas that require a thicker coating may experience a reduction in environmental life, or areas that require a thinner coating may experience a reduction in mechanical life. At least one type of turbine blade with a thin aluminum coating in the airfoil is known to have experienced airfoil internal oxidation. However, due to high shank stresses and technical challenges relating to the size of the blade, the internal coating is targeted to meet the shank requirement (less than 0.0254 mm or 0.001 inch coating thickness) and is the same throughout the internal cavities.
There is at least one known pack coating process, described in patent application Publication No. U.S. 2003/0211242, published Nov. 13, 2003, that coats an entire internal passage with a single coating thickness. However, small blades or other objects cannot be plumbed with vapor phase coating (VPC) to target a different coating thickness to different locations using this process.